Missile and launcher system



May 17, 1966 w. E. HERRMANN 3,251,301

MI'SSILE AND LAUNCHER SYSTEM Filed Sept. l2, 1962 6 Sheets-Sheet 1 Flsl INVENTOR WERNER E. HERRMANN Ageni` May 17, 1966 w. E. HERRMANN 3,251,301

MISSILE AND LANCHER SYSTEM Filed Sept. l2, 1962 6 Sheets-Sheet 2 /127 22x 24 FIG. 2

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WERNER E HERRMANN o o BY May 17, 1966 w. E. HERRMANN 3,251,301

MISSILE AND LAUNCHER SYSTEM Filed Sept. 12, 1962 6 Sheets-Sheet 5 `Agent May 17, 1966 w. E. Hr-:RRMANN MISSILE AND LAUNCHER SYSTEM 6 Sheets-Sheet 4 Filed Sept. l2, 1962 Nm Om vw INVETOR. WERNER E. HERRMANN Agent May 17, 1966 w. E. HERRMANN 3,251,301

MISSILE AND LAUNCHER SYSTEM Filed sept. 12, 1962 6 Sheets-Sheet 5 INVENTOR WERNER E.HERRMANN May 17, 1966 w. E. HERRMANN MISSILE AND LAUNCHER SYSTEM 6 Sheets-Sheet 6 Filed Sept. l2, 1962 INVENTOR. WERNER E. HERRMANN United States Patent() 3,251,301 MISSILE AND LAUNCHER SYSTEM Werner E. Herrmann, Northridge, Calif., assignor to Lockheed Aircraft Corporation, Burbank, Calif. Filed Sept. 12, 1962, Ser. No. 223,100 7 Claims. (Cl. 1112-49) This invention is directed to a missile and launch system and more particularly to an improved multi-purpose missile and compace launching system therefor.

The invention described herein provides various improvements over a basic invention which is a subject of patent application Ser. No. 51,621, led August 24, 1960,

now Patent No. 3,124,072 and entitled Missile Propulsion.

Combat troops in battlefield oper-ations have been universally presented with the problem of accurately and quickly obtaining intelligence information concerning enemy activities in support of their own tactical operations. The acquisition of such information in a readily useable form with minimum expense and personnel training have been particular problems. Additionally, the

` ability to achieve such information without substantial risk to personnel has been a virtually insoluble problem.

The existence of equipment of the character described in the referenced patent application provided a major step forward, particularly in the art of battlefield intelligence gathering. However, various aspects of such devices provided major problems for ultimate utility. For example, the providing of troops utilizing such equipment with facilities for stowing that equipment and with adequate launching capabilities, together with the development of certain structural means within missile itself have been points of concern. The maintenance of flight stability and the ability to recover the payload without destructive effect thereto have also presented difficulties. It is an object of the present invention to provide a system which is capable of providing small combat units with a self-suflicient, rapid and economical means for obtaining reliable battlefield information from enemy occupied areas.

Another object is to provide a missile and launch system capable of facilitating the acquisition of immediate day and night photographic intelligence to detect current enemy activities, supplement map data and substantiate the findings of other intelligence systems.

Yet another object is to provide a system capable of locating potential targets, supporting re adjustment and assessing damage in target areas in a rapid and relatively inexpensive manner.

A further object is to provide a missile and launch system having operational simplicity, relatively light weight and including all necessary components to complete the assigned mission. v

A still further object is to provide an intelligence gathering missile and launch system including means for recovery which substantially prevents damage to payload upon impact.

These and other related objects will become apparent from the following'description of certain preferred embodiments shown in the accompanying drawings in which:

FIGURE 1 is an elevational view of a missile and its launchingequipment stowed for transport;

FIGURE 2 is an exploded view of the launching equipment, the missile remaining in its launcher;

FIGURE 3 is a View of the missile and launching equipment assembled for launching the missile at an angle;

FIGURE 4 illustrates the launcher assembled with the missile for vertical launch;

FIGURES 5a, 5b and 5c show longitudinal cutaway views of various portions of the rocket assembly in a pre- 'ferred embodiment;

ICC

-FIGURE 6 is a section taken along line 6 6 of FIG- URE Sa to illustrate the missile separating mechanism;

FIGURE 7 is a perspective view of the missile fins in assembled condition; and

FIGURE 8 is a cutaway of the recovery balloon in deployed condition.

The missile and launcherV of this invention generally comprise a self-contained launching means and a rocket stowed within such means for transportation purposes, the missile being adapted to be launched therefrom. The rocket including propulsion and recoverable sections adapted for separation at predetermined operational instant, and recovery means adapted to surround the recoverable portion preliminary to its recovery impact.

More specifically, FIGURE 1 illustrates the preferred embodiment of the carrying and launching apparat-us of the invention, the missile and the fin assembly being enclosed in their stowed positions within the launcher. The launching means is illustrated more specifically in the exploded view of FIGURE 2.

The complete assembly in its transportable condition is indicated in FIGURE 1 by the numeral 10. The as- -sembly 10 includes a launching tube 11 (shown more for the gripping and flexing outward of the free end of the tabs 13. The smallest internal diameter of the launching tube 11 is dependent upon the external diameter of the rocket stowable therein and launchable therefrom. It is .a requirement that the tube be of sufficient diameter to facilitate free longitudinal movement of the rocket therethrough, but that it be suiiiciently small to prevent excessive lateral movement of the rocket during transport phases. The launching tube 11 may be recessed at its upper and lower extremities 15 and 16 to facilitate the adaptation of connecting or cover means. The tube is preferably fabricated from a light weight, but sturdy material such as aluminum or plastic.

A mid case 18 is usually provided and adapted for juxtapositioning with respect to recessed portion 1S of tube -11. The mid case 18 has an ex-ternal configuration similar to that of the launching tube 11 and normally includes recessed portions 19 and 20 at its extremities. An upper case 22 adapted to be juxtapositioned against recessed portion 20 of the mid case 18 is preferably of a square cross sectional configuration and includes an internal area suciently large .to accept a fin and collar assembly in knocked down condition, to be described later. The upper case 22 usually includes recessed portions 23 and 24 at its lower and upper extremities, respectively Case caps 26 and 27 are provided to cover the respective lower and upper ends of the case or carrying assembly throughout transport and storage phases.

When each of the case sections and the caps are juxtaposed longitudinally with respect to one another i-n the manner illustrated in FIGURE 1, attachment means indicated as a, Ab, c and d, of a conventional nature such as tape, metal bands, etc. are disposed about the juxtaposed recessed portions to prevent separation of the sections during transportation. It is, of course, possible to facilitate such connections without the described recesses. Additionally, conventional lock means upon the various case portions may be provided to prevent separation under conditions of stress where such are deemed necessary.

A plurality of legs or stanchons 29 are provided lfor the support of the launching tube during launching phases of operation. The legs 29 are preferably of substantially triangular cross sections configuration and are usually of a length approximating that of the launching tube 11 plus the mid case 18. The legs 29, when stowed for transportation purposes (FIGURE l) are normally positioned externally of the launching tube 11 and the mid case 18 so as to cover the perforations 12 and the tabs tl3 and prevent the entrance into the launching tube of foreign matter. The legs 29, when so stowed, are normally over-wrapped with tape, as at e and f, or otherwise retained by conventional means. Longitudinal tapes covering open regions between the legs 29 and the launch tube 11 may also be provided. Pull-cords are normally provided with all tapes to facilitate easy disassembly from. the FIGURE 1 condition and rapid and convenient launching set-up.

Prior to the assembly of the described launching and transport apparatus, a missile (described below) is inserted therein and a fin and collar assembly in disassembled condition is inserted into the upper ca-se 22.

In preparation for launching procedures, the assembly being initially in the stowed condition of FIGURE 1, disassembly is first accomplished by removal of the connecting means, e.g., by pulling pull cords in the tapes, thereby allowing the upper ca-se 22, the mid case 18, the case caps 24 and 26 and the legs 29 to be removed. In the event that a launch at an angled trajectory is to be accomplished a pair of adjacent tabs 13 are bent outward and legs 29 are inserted thereover in the manner illustrated in FIGURE 3. The trajectory angle may be varied by adjusting the angular spread between the launching case 11 and the legs 29.

A vertical launch may also be accomplished by afxing all 4 of the legs 29 over the respective tabs 13 in essenytially the manner illustrated in FIGURE 4.

This launching apparatus may be reused an indeterminate number of times.

The missile portion of this invention was described in a basic configuration in the above referenced co-pending patent application. The rocket of the present invention as illustrated in detail in FIGURES a, 5b and 5c is generally indicated by numeral 30 and includes a first rocket 31, a payload portion 32 and -a second rocket'33.

The rst rocket 31 is preferably rotatable and removably connected to the payload portion 32. The payload portion 32 is usually similarly rotatable and removably connected to the second rocket 33 at its opposite extremity.

The major portion of first rocket 31 is constructed in a basically conventional manner. An elongated case 35 includes an insulation 36 and the combustion inhibitor 37 concentrically disposed therein and a conventional solid propellent grain 38 disposed intermediately of the combustion inhibitor. An igniter 39 of conventional configuration is disposed internally of a longitudinal channel 40 intermediate of grain 38. A pluralityof nozzles 41 (2 being shown in the present configuration) are provided at the after end of the rocket 31 and are directed outward and rearward. The nozzles 41 are preferably canted at a predetermined angle to impart spin to rocket 31 during its operation. Additionally the nozzles 41 are some-times scrafed at the locations indicated as 42 to prevent interference between ltheir outer extremities on lthe launching structure. Rigidly attached to the after extremity of the rocket 31 upon the longitudinal axis thereof is a bearing extension 43 having a peripheral ange 44 aiixed thereto. Surrounding the outer surface of the extension 43 and disposed over the forward surface of `the peripheral flange 44 is a bearing material 45 such as the materials commonly and commercially known as nylon, Teflon, or other similar materials having structural and lubricating qualities. Surrounding the bearing extension 43 and the bearing material 45 in bearing relation is an outer bearing journal 46. Rigidly attached to and extending rearwardly from the bearing journal 45 is a housing 47 including an inner cavity 48. An annular channel 49 is dened in the inner cavity 48 of the housing 47 adjacent the after extremity of the housing 4'7, the channel 49 being adapted to accept locking means for locking the rocket 31 and the payload portion 32 in longitudinally disposed relation.

The payload portion 32 generally includes a forward section 5t), which usually carries the payload, anda vrearward section 51, which normally includes the recovery means. Although the payload is not a portion of this invention, a camera mechanism 52 is shown to illustrate the adaptability of the present invention to photographic reconnaissance purposes. The lforward section 50, when adapted to photographic purposes as herein illustrated, includes a protective shield 53 over the lens of the camera 52, the shield 53 having a perforation 54 through its extremity to facilitate accomplishment of photographic procedures therethrough. The shield 53 is attached to a housing 55 and is therewith adapted for insertion within inner cavity 48 of housing 47.

The housing 55 includes therein an annular groove 56 and a series of radially extending perforations 57, best shown in FIGURE 6. A slotted lock ring 58 is retained in the annular groove 56 by means, for example, of an expansion type lock ring 59 and such that a series of lands 60 upon the outer periphery of the slotted ring 58 slidably engage the periphery of the annular groove 56. The land peripheries bearagainst a series of balls 61 disposed within radial perforations 57, thereby causing the yballs 61 to become engaged within the annular groove 49 of the housing 46 and preventing relative axial movement between the rocket 31 and the payload portion 32. Each of the lands 60 is provided with a ramp 60a on one side thereof to facilitate the movement of the balls 61 radially outward as the ring 58 is rotated by appropriate tool means, thereby accomplishing the locking. A squib means or other conventional actuator means 62 and detent Imeans 62a engageable with holes 63 (FIGURE 6) in lock ring 58 are provided for rotating said ring responsive to a predetermined signal. This rotation will be sucient to allow the balls 61 to fall into the regions between the ylands 60, thus disengaging the balls 61 from the annular groove 49 and allowing separation of the `rocket 31 from the payload portion 32. During the tiring sequence of the rocket 31 relative rotation between rocket 31 and payload portion 32 may be facili-tated at both the bearing`44 location and between the housings 46 and 55, the balls 61 acting as a bearing in the latter event.

Detent means 64 is also sometimes provided on one end of the forward section 5) to prevent removal therefrom of the lin assembly hereinafter described. The detent means may be of the illustrated spring loaded character or of other conventional construction, e.g., including retraction means responsive to a predetermined signal or condition.

The second rocket 33 and the rearward section 51 of the payload portion 32 m-ay be constructed in a manner similar to or identical with that explained with respect to rocket 31 and forward section 50 of the payload portion 52. Alternatively, it is sometimes not necessary that the journal bearing relationship be included. As illustrated, the rocket 33 includes a tail skirt 65 which is completely hollow in order to accept payload recovery means or other matter desired to be stored in that location. The rocket 33 is separable from the payload portion 32 in the same manner as described above with respect to the rocket 31 and the payload portion 312.

A fin assembly 66 includes a collar 67 and a plurality of fins 61S. This lassembly is shown in its disassembled condition in FIGURE 2, in its stowed position in FIG- U-RE 1 by broken lines and, for purposes of clarity, it is illustrated in perspective in FIGURE 7. The iin assembly is provided for flight stability purposes, particularly during the initial phase of launching and prior to ring of the retro-phase hereinafter described. The fin collar 67 is provided with the series of slots 69 through which the tins 68 are inserted after their removal from the upper case 22 and prior to launching.

As shown in FIGURE 5b, the forward and rearward sections 50 and 51 of the payload portion 32 are preferably intermediately connected thereof by a conventional U clamp 70 or similar'means to facilitate accessibility to the interior of the payload portion.

Referring now to FIGURES 5b and 8, the rearward section 51 usually has payload section recovery means stored therein. vAlthough parachutes and similar equipment may be readily utilized in this respect, the preferred recovery means is illustrated as being stowed in a removable nose section or shield 72. This recovery means, which may be referred to as a recovery balloon and which is illustrated in its deployed condition in FIGURE 8, is actuated by a conventional pressure means 73 responsive to a predetermined signal controlled by any conventional time delay, sequencing, or similar apparatus. The recovery balloon 74 is preferably fabricated from a iiexible and substantially non-stretc-hable material which is highly resistant to chatting, tearing, puncturing, etc. Materials such as plastic or rubber impregnated fabric may be uti- A applied. Continued expansion results in an ultimate coniiguration substantially as illustrated in FIGUR'E 8. It will be noted that therein the external periphery of the recovery balloon 74 is substantially ovoidal, a tapered channel region 75 axially intermediate of the balloon 74 being disposed about and extendin-g beyond both ends of the payload portion 32. The tapered characteristic of the channel 75 facilitates pressurized deployment of the recovery balloon 74 without substantial structural interference. A lip portion 76- interiorally of the channel 7'5 may be provided for clamping the recovery balloon 74 to the after end of the rearward payload section 51. `Although the recovery balloon 74 is illustrated as being generally ovoidal it is sometimes preferred that a spherical-cross-sec-tional configuration be utilized. It is also desirable upon occasion that the balloon not be fully inilated in order that the tendency of the recovered package to bounce upon ground contact will be dampened. The ends 76 and 77 of the recovery balloon 74 should extend beyond the ends of the payload portion 32 a distance sufiicient to insure that a landing upon either end of thel payload section will not result in a direct impact of the payload section with the landing terrain.

operationally, a launching of the missile for a penetration mision, i.e., including retro-phase operation, is .begun with the launching mechanism in the position essentially as illustrated in FIGURE 3. The rocket 30 is inserted into the launch tube 11 until such time as it bottoms on stop means (not shown) provided for that purpose in launch tube 111. At this juncture the first rocket 3K1 is protruding from the launch tube 11 and the major portion of the payload portion 32 and the second rocket 33 are disposed within the launch tube. The fins 68 are inserted within the slots 69 of the collar 67 and the lin `assembly is slipped over the iirst rocket 31 into the position illustrated in FIGURE 3. The ignition of propellent grain 38 is accomplished by igniter 39 such that thrust and a rotational moment are imparted to 6 rocket 31. The thrust is transmitted through the extension 43 and the bearing 44 to the payload portion 32 via the detent balls 61. It is then transmitted into the rocket 33 via similar balls in the after detent and releaseV portion. Hence, forward movement of the rocket 31 results in the ultimate movement of the entire rocket assembly. Since the interior of the n collar 67 is of diameter greater than the external diameter of the payload portion 32 Iand the Iforward portion of the rocket 33, those portions of the rocket 30 move through the n assembly 65 until such time as the fin assembly is engaged by the stop means 71 upon the after end of the rocket 33, the iin asserri'bly 66 being then carried with the balance of the rocket 30 throughout this initial boost phase and trajectory. This launch or boost phase of rocket 30 is stabilized both by the rotation of the rocket 31 and by the laction of the ltins 66. At a predetermined instant while the rocket 31 is still thrusting, the actuator means 62 rotates the slotted lock ring 5-8 such that the balls 61 slip inward into the regions between the lands' 60, thereby allowing the rocket 3'1 to move from engagement with the payload portion 32. When the rocket 31 burns out it falls away and the balance of the rocket continues on its trajectory. At a predetermined instant the rocket 33, which is in actuality a retro-rocket, is tired. The firing of the retro-rocket 33 causes the assembly to decelerate. `During this phase of operation the iiight characteristics are stabilized by the presence of the tins 68. Eventually, the thrust of the retro-rocket 33 overcomes the inertia of the assembly, stopping it in m-idair and accelerating it in the reverse direction. This acceleration causes the retro-rocket 33 and the payload portion 32 to move through the iin assembly 66 until lsuch time as the fin assembly drops from the noW rear- Ward end of the payload portion 32 in some applications, or until the fins engage the detent means 64. In the latter event the iins are thereafter carried in this position for stability purposes or dropped in response to a predetermined signal or condition causing retraction of detent means 64. The drag forces on the ns assist in their ultimate removal. Simultaneously, the canted nozzles of the retro-rocket 33 cause its rotation with respect to the payload portion 32 upon the interconnecting ball `detent mechanism, thereby stabilizing the retro flight. Also at a predetermined instant the camera mechanism 52 becomes operational and photographs of the terrain thereunder are taken.

Prior to burn-out of the retro-rocket 33 its separation mechanism is actuated and the retro-rocket moves away from the payload portion 32, the payload portion 32 continuing on its new trajectory.

At a predetermined instant during the descent of the payload portion 32 the pressure means 73 is actuated and the recovery balloon 74 is deployed. Upon landing of the payload portion 32 with the recovery balloon 74 so deployed this segment of the rocket 30 may be recovered and the intelligence information in the payload portion utilized in any desirable manner.

When the rocket is to be launched vertically, it is sometimes expedient to use a single rocket, rather than one on either end of the payload portion. In such event either of the rockets 31 or 33 may be utilized.

It will be apparent from the description that the payload portion 32 of this rocket assembly will be reuseable when recovered. It will also be apparent that the basic invention described herein may be utilized for a variety of end uses with various types of payloads.

Whilespecic embodiments of the invention have been shown and described it should be'understood that certain alterations, modifications and substitutions may be made to the instant disclosure without departing from the spirit and scope of the invention as defined by the appended claims.

I claim:

1. A missile comprising:

a payload portion having a iirst and a second end,

a dirst rocket being connected to the first end of the payload portion and being used for propelling the missile along a predetermined first trajectory,

`a second rocket being connected to the payload portion `second end and used for propelling the missile along a predetermined se'cond trajectory,

=both the first and second rockets being rotatable rwith respect to and removable from the payload portion,

means on the first and second rockets for rotating the rockets relative to the payload portion, and means for stabilizing the rocket in both the first and second trajectory and being reciprocally mounted on the payload portion and movable into proximity of the second rocket when the rst rocket is fired and movable into proximity of the first rocket when the second rocket is fired.

2. The missile of claim 1 wherein (a) said payload portion is provided with recovery meansactuable to' substantially surrounded said payload portion following separation of said rockets therefrom.

3. A missile comprising:

a payload portion having a first and second end,

a first missile-propelling rocket rotatable and removably connected in pulling relation to the payload portion first end,

a second missile-propelling rocket rotatably and removably connected in pulling relation to the payload portion second end,

means being disposed on the first and second ends of the payload portion for separating the iirst and second rockets,

the rst and second rockets each including an annular groove,

the separation means including a plurality of balls retained in engagement with the rocket grooves by a ring means, the balls being movable radially inward to disengage the payload portion and the rockets and responsive to rotation of the ring means, and a means for rotating the ring means, and

a recovery means being stored in the payload section for deployment after yseparation of the second rocket.

4. The missile of claim 3 wherein said recovery means is a balloon stored in one end of and within the major diameter of said payload section and for expanding into a shape radially surrounding and longitudinally overextending the ends of said payload section.

5. The missile of claim 4 wherein said balloon is of substantially ovoidal external configuration when employed.

6. The missile of claim 4 wherein said balloon is retained adjacent one end of said payload portion and includes means defining a longitudinal channel when deployed within which said payload portion is disposed, said channel means being flared outward from said attachment to the opposite channel means extremity.

7. A missile comprising (a) a first pull rocket adapted to initiate missile flight and to spin throughout its operation,

(b) a payload portion having a first end connected to said first rocket,

(c) bearing means between said first rocket and said payload portion to facilitate relative rotation,

(d) separation means mounted in the payload portion responsive to a predetermined first condition for separating said first rocket and said payload portion,

(e) a second rocket connected to a second end of said payload portion and adapted to substantially reverse missile flight direction upon attainment of a predetermined second condition,

(f) fin means positioned upon said second rocket during an initial portion of said missile flight and separable from said missile during said reverse flight,

(g) bearing means between said second rocket and said payload portion to facilitate relative rotation,

(h) a second separation means mounted in the payload portion responsive to a predetermined second condition for separating said second rocket and said payload portion, and

(i) balloon means stowed in one end of and within the major diameter of said payload portion for surrounding said payload portion upon attainment of apredetermined third condition.

References Cited by the Examiner UNITED STATES PATENTS 2,344,957 3/ 1944 Anzalone l02-34.1 2,452,783 11/ 1948 Nebesar 244-138 2,623,465 12/ 1952 Jasse 102-49 2,655,105 10/1953 Hansche 102-49 2,656,135 10/1953 Barker et al 102-50 X 2,804,823 9/ 1957 Jablansky 102-49 2,835,199 5/1958 Stanly 102-50 2,959,129 11/1960 Warren 102-49 2,972,933 2/ 1961 Guthrie et al 89-1.7 2,981,187 4/1961 Riodan et al 102-49 3,015,992 1/1962 Ransom 89--1.7 3,053,476 9/1962 Mohar 244-138 X 3,067,682 12/ 1962 Feldmann et a1. 102-49 3,124,072 3/ 1964 Hermann 102-49 BENJAMIN A. vBORCHELT, Primary Examiner.

SAMUEL FEINBERG, Examiner. 

1. A MISSILE COMPRISING: A PAYLOAD PORTION HAVING A FIRST AND A SECOND END, A FIRST ROCKET BEING CONNECTED TO THE FIRST END OF THE PAYLOAD PORTION AND BEING USED FOR PROPELLING THE MISSILE ALONG A PREDETERMINED FIRST TRAJECTORY, A SECOND ROCKET BEING CONNECTED TO THE PAYLOAD PORTION SECOND END AND USED FOR PROPELLING THE MISSILE ALONG A PREDETERMINED SECOND TRAJECTORY, BOTH THE FIRST AND SECOND ROCKETS BEING ROTATABLE WITH RESPECT TO AND REMOVABLE FROM THE PAYLOAD PORTION, MEANS ON THE FIRST AND SECOND ROCKETS FOR ROTATING THE ROCKETS RELATIVE TO THE PAYLOAD PORTION, AND MEANS FOR STABILIZING THE ROCKET IN BOTH THE FIRST AND SECOND TRAJECTORY AND BEING RECIPROCALLY MOUNTED ON THE PAYLOAD PORTION AND MOVABLE INTO PROXIMITY OF THE SECOND ROCKET WHEN THE FIRST ROCKET IS FIRED AND MOVABLE INTO PROXIMITY OF THE FIRST ROCKET WHEN THE SECOND ROCKET IS FIRED. 